CN103161513B - The improved nozzle vane for gas-turbine unit - Google Patents
The improved nozzle vane for gas-turbine unit Download PDFInfo
- Publication number
- CN103161513B CN103161513B CN201210545238.3A CN201210545238A CN103161513B CN 103161513 B CN103161513 B CN 103161513B CN 201210545238 A CN201210545238 A CN 201210545238A CN 103161513 B CN103161513 B CN 103161513B
- Authority
- CN
- China
- Prior art keywords
- nozzle vane
- impinging
- impinging cooling
- wall surface
- nozzle
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired - Fee Related
Links
- 238000001816 cooling Methods 0.000 claims abstract description 67
- 239000002826 coolant Substances 0.000 claims abstract description 35
- 239000012530 fluid Substances 0.000 claims abstract description 27
- 238000012546 transfer Methods 0.000 claims description 25
- 238000003780 insertion Methods 0.000 claims description 23
- 230000037431 insertion Effects 0.000 claims description 23
- 238000011144 upstream manufacturing Methods 0.000 claims description 6
- 230000005855 radiation Effects 0.000 claims description 4
- 238000003491 array Methods 0.000 claims description 3
- 230000015556 catabolic process Effects 0.000 claims description 2
- 238000006731 degradation reaction Methods 0.000 claims description 2
- 239000000203 mixture Substances 0.000 claims description 2
- 239000007921 spray Substances 0.000 claims description 2
- 238000005452 bending Methods 0.000 claims 1
- 238000013461 design Methods 0.000 description 7
- 230000000694 effects Effects 0.000 description 6
- 230000008859 change Effects 0.000 description 5
- 239000000567 combustion gas Substances 0.000 description 5
- 239000007787 solid Substances 0.000 description 5
- 238000002485 combustion reaction Methods 0.000 description 4
- 239000012809 cooling fluid Substances 0.000 description 4
- 241000194386 Coelosis Species 0.000 description 3
- 238000010276 construction Methods 0.000 description 3
- 239000007789 gas Substances 0.000 description 3
- 239000002184 metal Substances 0.000 description 3
- 238000009826 distribution Methods 0.000 description 2
- 239000000463 material Substances 0.000 description 2
- 238000000034 method Methods 0.000 description 2
- 238000006386 neutralization reaction Methods 0.000 description 2
- 238000005086 pumping Methods 0.000 description 2
- 230000015572 biosynthetic process Effects 0.000 description 1
- 230000000903 blocking effect Effects 0.000 description 1
- 238000004891 communication Methods 0.000 description 1
- 230000006835 compression Effects 0.000 description 1
- 238000007906 compression Methods 0.000 description 1
- 230000001419 dependent effect Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000005611 electricity Effects 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 239000004744 fabric Substances 0.000 description 1
- 230000003116 impacting effect Effects 0.000 description 1
- 230000006872 improvement Effects 0.000 description 1
- 238000002347 injection Methods 0.000 description 1
- 239000007924 injection Substances 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 230000007246 mechanism Effects 0.000 description 1
- 238000004080 punching Methods 0.000 description 1
- 230000009467 reduction Effects 0.000 description 1
- 238000007634 remodeling Methods 0.000 description 1
- 238000010025 steaming Methods 0.000 description 1
- 239000013589 supplement Substances 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
- F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
- F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/06—Fluid supply conduits to nozzles or the like
- F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/10—Two-dimensional
- F05D2250/13—Two-dimensional trapezoidal
- F05D2250/131—Two-dimensional trapezoidal polygonal
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The present invention provides a kind of improved nozzle vane for gas-turbine unit, and this nozzle vane includes：Blade wall, this blade wall has inner wall surface and outer wall surface, and described wall surface is spaced apart from each other, to limit the multiple fluid passages for cooling medium；Discrete chamber, described discrete chamber is formed by the inside wall member being arranged in the fluid passage between inner wall surface and outer wall surface and for cooling medium；Multiple impinging cooling sleeves, in the discrete chamber that the plurality of impinging cooling sleeve is arranged in by inner wall surface and outer wall surface and is limited by internal wall member；And multiple non-circular, the such as zigzag in each of impinging cooling sleeve impinging cooling sleeve, opening, the size and number of described opening be enough to accommodate the flowing of cooling medium.
Description
Technical field
The present invention relates to be for example used for electric those gas-turbine units generating electricity with aircraft engine, and more specifically
Ground, relates to nozzle chambers and blade design using having improved convection current and impinging cooling ability and comes cooling nozzles and correlation
The system of the blade of connection.The present invention can also be used to other engine components, for example cover, bucket leaf/blade (buckets), combustion
Burner and being used together from different airfoils and nozzle inner wall and nozzle outer sidewall.
Background technology
For many years, to protect the steam of gas-turbine unit during long-time operation using various cooling bodies
Path components, particularly often meet with the first order jet nozzle of temperature highest aerofluxuss.Most of closed-loop paths cooling system includes
The multiple nozzle vane sections extending between the medial wall of nozzle and lateral wall.Typically, blade includes chamber, described chamber and lateral wall
It is in fluid communication with the room in medial wall, to accommodate the flowing for cooling down the cooling medium in outer wall and the closed-loop path of inwall.
Cooling medium is supplied in the pumping chamber in outer wall, logical by the different room of impact limited opening and flowing for being assigned to
Road, thus allow coolant to flow in the outer wall surface of blade.Used impinging cooling medium then flows into leading edge and footpath
To the back cavity extending through blade.
In the past, steam is used as the cooling medium of certain form of nozzle vane.Although under nominal operating temperature
(nominal operating temperatures), the heat capacity ratio air that steam has is high, but is used for turbine blade
Show certain pinodynainic inefficiency with the steam Cooling Design of other engine components.For example it is necessary to be maintained at steam
Inside closed-loop path, to avoid mixing with thermal current.Therefore, some parts in hot gas path can not be as closed-loop path inside
Steam the same be efficiently cooled, reason is that the relatively thin structure of the trailing edge of such as nozzle vane prevents blade some
Partial effective steam cooling.
Another kind of known method for cooling combustion turbine machine electromotor wheel blade and nozzle includes cooling down using high pressure
The part supply of air, described high pressure cooling air is generally by the intergrade of such as GTC or the inside of final level
Source provides.Typically, nozzle neutralization uses gaseous film control to provide the expectation of blade around a series of internal flow passageway of nozzle
Supplement cooling, and exterior line supplies compressed air to nozzle, and described compressed air is finally disposed to combustion gas turbine
Thermal current in.
Most of current combustion gas turbine also relies on some form of impact heat transfer, with by putting with respect to blade surface
Put one group of circular port (a bank of round holes) and introduce relatively high speed fluid (steaming directly with respect to the surface of solids
Vapour or air) spray cooling nozzles blade.The speed of cooling fluid is higher, molecule trend towards with surface keep contact and
The time of heat exchange is longer.For this reason, impinging cooling fluid injection is generally and metal surface is introduced vertically into air, so that phase
Incidence rate for described surface maximizes.In some recent design, impinging air cooling is used with steam combination, with
Reduce only by steam can not the effective specific part of the nozzle vane of cooling operation temperature.However, virtually all of only make
With air, the impinging cooling system for combustion gas turbine is dependent on specified quantity in blade for the circular port and arrangement to hold
Receive flow of fluid.
Although the higher biography of level of relative can be realized in first order nozzle using the impinging cooling with circular open
Heat, but once Contact -impact occurs, then fluid molecule trends towards being moved parallel to and fluid velocity with solid blade surface
Substantially become to reduce, less molecule is contacted with the surface of solids, ultimately result in heat transfer and reduce.Due to from adjacent circular impact
The fluid that hole enters can collide, mix and finally reduce coolant handling capacity (coolant throughput), cooling stream
Body speed also becomes much lower.Similarly, local pressure declines and trends towards redirecting flow of fluid, thus further dropping
Low fluid velocity.Due to always related to circular impact opening phenomenon, the heat transfer in this nozzle vane degrades and (is referred to as " horizontal
Stream effect ") so that heat transfer level is reduced.
It has therefore been discovered that by the air-flow character being obtained, use to cool down purpose compressed air and/
Or steam pays the cost that thermodynamic efficiency reduces to a certain extent using circular impact hole.Coolant and blade surface
Between heat transfer amount, with coolant, when it clashes into the surface that is cooled and turns next to the table being Yu being cooled
Speed when face is parallel is directly proportional.Therefore, hot surface to be cooled exists the heat transfer of the difference/change of discrete groups
Coefficient.Highest heat transfer is realized at just to impact opening, but reduces with the coolant velocity leaving hole and becomes relatively low.Due to coming
Coolant from adjacent holes is mixed with the coolant phase from circular impact hole, therefore from adjacent circular hole crossing current mutual
Effect also makes cooling effect reduce, so that its speed reduces and so that heat transfer potential is declined.
Therefore, however it remains to the notable needs determining such method：So that for blade cooling compressed air or
The heat transfer potential of steam maximizes, and thus holding member temperature under strict operation requires.Also exist and make coolant speed
While degree improves, so that the flowing in the maximum area of the blade surface being cooled more uniformly is needed, thus provide excellent
Elegant overall heat transfer efficiency.
As described above, solving the problems, such as that the current state of the technical field of impinging cooling almost relies only on circular impact hole to produce
Raw desired cooling effect.For example, see U.S.No.6,468,031 (describe to make the inside of airfoil using circular impact hole
The nozzle that heat transfer on face increases).Similarly, EP1247940A1 describes the use with the circular impact hole of variable-diameter,
To prevent blocking without making the heat transfer between coolant and nozzle surface reduce.
Content of the invention
The present invention provides a kind of improved nozzle vane for combustion gas turbine, and this nozzle vane uses dramatically different
Impact opening and the air flow being obtained.As described above, the present invention can also combination can have benefited from impinging cooling its
Its combustion turbine engine components uses, including cover, bucket leaf and burner and other types of airfoil and or even nozzle
Medial wall and nozzle outer sidewall.
In the exemplary embodiment being related to nozzle vane, the present invention includes：(1) blade wall, this blade wall has interior table
Face and outer surface, wall surface is spaced apart from each other, and is led to the fluid limiting for cooling medium (typically compression environment air)
Road；(2) multiple discrete chambeies, the plurality of discrete chamber is by being arranged between inner wall surface and outer wall surface and be used for cooling down
Inside wall member in blade in the fluid passage of medium is formed；(3) multiple impinging cooling sleeves, the plurality of impinging cooling
Sleeve is arranged in the discrete chamber being limited by inner wall surface and outer wall surface and internal wall member；And (4) impinging cooling
Multiple non-circular impact opening in each of sleeve impinging cooling sleeve, the size and number of described opening be enough to accommodate
The flowing of high pressure cooling medium.
In embodiment described in this manual, each of multiple impact openings impact opening is all formed as small-sized
Axial fluid path, described small-sized axial fluid path has the radial direction battle array of the protuberance from the radiation setting of the center of each opening
Row.Integrate, by way of example, described protuberance is formed from the radiation of the center of described opening, substantially homogeneous stream
The radial arrays (for example, in star form) of body path.Described protuberance can take various forms, including equidistant or alternately set
The protuberance put or there is different size and the protuberance of different distance separated from one another.
As discussed below, contrast circular open, non-circular in the impinging cooling sleeve described in this specification opens
Mouth provides mixing and the heat transfer potential of much more efficient.I.e. it has now been found that non-circular impact opening makes cooling fluid speed
Degree improves and crossing current effect (cross flow) is minimized, and thus increases the biography between cooling fluid and blade surface
Heat.New geometrical pattern design also creates rotational field (vortex field) at impact opening neutralization is around impact opening,
It provides more efficient mixing and improved air velocity distribution at each opening.Due to the improvement of opening make mixing and
The higher characteristic of speed, the discrete fluid passage radiating from the center of each opening (for example, there is the opening of star configuration)
This characteristic significantly improves the heat-transfer capability of existing nozzle design.
Additionally, be economically feasible using new impact opening, reason is that the formation of new impact opening does not need
Any significant design or operation change are carried out to the other parts of gas-turbine unit.Therefore, the present invention only passes through minimum
Extra manufacturing cost and complexity and improve various types of nozzles, blade, cover, bucket leaf, the rushing of burner and airfoil
Hit heat transfer potential.The raising of heat transfer also improves the life expectancy of critical engine parts, is especially exposed to combustion gas turbine
In high gas path temperature those critical engine parts.Specifically, prominent from center radiation setting using having
Multiple axial fluid paths of the radial arrays in portion finally make metal temperature relatively low and crucial hot gas path part longevity
Life is longer.According to concrete application, this heat transfer raising reduces engine coolant requirement, and has finally improved the total of electromotor
Body performance and thermodynamic efficiency.
Can be implemented in various parts according to the use that the multiaspect of the present invention impacts opening, such as nozzle, the leaf that struggles against, cover,
Burner and housing.Required using application, target component and cooled flow according to specifically final, star-shaped openings specifically several
What size can also slightly change.Therefore, invention contemplates impacting the variously-shaped, including for example of opening using non-circular
" herringbone " (" chevron ") class type hole or other symmetric construction.When be applied to primary stage nozzle, first order bucket leaf,
When combustion liner and/or case temperature control device, this is designed to improve the heat-transfer capability of those parts.
By the more detailed description to currently preferred exemplary embodiment below in conjunction with accompanying drawing, the present invention these
And other objects and advantages will be more complete understanding of and be understood.
Brief description
Fig. 1 is the schematic partial perspective view of the nozzle vane with cooling insertion sleeve, described cooling insertion sleeve cloth
Put in the blade cavity with the impact opening implementing the present invention；
Fig. 2 is the further perspective view with the new exemplary insertion sleeve of impact opening；
Fig. 3 is the zoomed-in view of some selected impact opening shown in the embodiment of Fig. 1；And
Fig. 4 is the schematical vertical cross section of another insertion sleeve implementing the present invention.
Specific embodiment
As described above, improve according to the impact opening of the present invention conducting heat and improve cooling air and/or steam
Effectively use, to reduce the temperature of nozzle and other engine components.Contrast has the circular open of the cross-sectional area being consistent, limit
That has determined that multiple Geometry edges of opening have is sized and configured to so as to impact jet velocity higher and so that heat transfer is increased
Plus.The handling capacity (throughput) bigger using permission of this multiaspect impact opening and the net reduction of metal temperature.Also find,
Star configuration have produce vortex flow field ability, this vortex flow field energy enough provide cooling air before acting on blade more
Efficient mixing.More efficient mixing allows the higher distribution of the speed of cooling air, and therefore allow fluid and surface it
Between higher heat transfer.
The general type of exemplary insertion sleeve is shown in Fig. 1 to Fig. 3.Fig. 1 shows the sleeve for leading edge chamber, and
Fig. 2 and Fig. 3 shows the exemplary sleeve in the chamber 17 being used in particular in Fig. 1.In the embodiment in figure 1, nozzle vane 10 includes
Multiple collection coelosis (chamber being wholely set, integral cavities) 11,12,13,14,15,16 and 17, in multiple collection coelosis
Each collection coelosis there is insertion sleeve, insertion sleeve 18,19,20,21,22 and 23 is in have many surface currents hereinafter described
The general type of body path, the hollow and sleeve of general rectangular.Described sleeve is configured to the shape with the particular cavity of receiving sleeve
Shape closely corresponds to, and the both sides of sleeve have multiple impinging cooling openings along the sleeve part with adjacent chamber walls.
In the leading edge chamber 30 of Fig. 1, the forward edge of insertion sleeve 31 has curved configuration, and the shape of side wall generally corresponds to
Shape in the side wall in multiaspect (multi-faceted) chamber 30.The side wall of insertion sleeve is many along the part inclusion of whole length
Individual impact opening.As illustrated, impact opening has multiaspect, i.e. zigzag configuration (is also shown as the impact opening 40 amplifying).Insert
The rear side 32 entering sleeve 31 does not include any impact opening.Similarly, in back cavity 13,12 and 11, insertion sleeve 19 and 18
Side wall only has impact opening along a part for length, and the antetheca of insertion sleeve 19 and 18 and rear wall are by solid non-perforated material
(solid non-perforated material) is formed.Many surface constructions shown in Fig. 1 are only opening of being conceived of the present invention
One example of mouth type.Can be using other geometric constructions with all even non-uniform zigzag (serrations), such as
Multi-stylus star, herringbone type hole, " spike shape " opening etc..
Sleeve in chamber 30,17,16,15,14,13 and 12 is spaced apart with the wall in chamber, so that cooling medium (for example, is pressed
Contracting air) into and through impact opening, thus acting on the interior wall surface in chamber, thus cooling wall in the manner described above
Surface.It is travel downwardly from the upstream end in chamber with impinging cooling agent, crossing current degradation trends towards increasing and generally will result in relatively low
Heat transfer.However, new impinging cooling opening shows relatively low pressure drop in the length of blade, provide more efficient simultaneously
Cooling.
As Fig. 1 also shown in, the impinging cooling opening in upstream portion that insertion sleeve includes be arranged in sleeve.Other
Downstream part is basic atresia and does not include hole, but by reducing in the area between insertion sleeve and chamber interior wall
Coolant flow area and play the effect of blockage mechanism.See sleeve gap 50.Therefore, using the design of new impact opening
So that mixing improves, when air makes impinging air jet velocity improve when opening moves to target surface and makes along blade
Whole length heat transfer coefficient finally more preferable.New opening also reduces undesirable rear impact coolant crossing current.
Insertion sleeve 23 shown in Fig. 1 and Fig. 2 includes the elongate sleeve with open lower end, and this open lower end has side
Edge flange, for connecting to the opening of respective chamber, the such as chamber 17 in Fig. 1.The side wall 50,51 of sleeve 23 includes multiple respectively
Zigzag impinging cooling opening 52,53, is shown specifically at 40 in the Fig. 3 such as amplifying.Impinging cooling hole 52,53 is along sleeve
First upstream sleeve part 55 and 56 is defined, for making cooling medium flow into sleeve and inner vanes wall surface to be cooled
Between space in.Sawtooth on opening also creates rotational field (vortex field), and this rotational field makes interface outside impact
Interior mixing is improved and hence in so that the impact jet velocity acting on cavity wall is higher, thus then producing higher
The more effective use of heat transfer coefficient, relatively low metallic surface temperature and obtainable cooling air.The downstream sleeve part of sleeve
57,58 are divided not have impact opening.On the contrary, downstream part reduces the coolant flow in chamber 17 by limiting fluid flowing passage
Dynamic area, described fluid flowing passage has benefited from backlash by the space of the first impact opening being defined as in neighbouring sleeve
Hit cooled flow.
As shown in Figure 2, the scope (extent) being provided with the part of sleeve of impact opening 52 and 53 depends on inserting
Enter sleeve lateral wall towards be airfoil on the pressure side or suction side.Although if being considered as to need or expect, punching
Hit scope on every side for the hole can change, but the scope of the impact on the pressure side gone up of sleeve is preferably more than suction side
's.
With reference to Fig. 4, insertion sleeve 60 is arranged in blade cavity 19.The circumferential profile of insertion sleeve 60 equally along/follow
The profile in chamber and have on side wall 63,64 star impact opening 61,62.Coolant (for example, compressed air) is from pumping chamber
Flow in insertion sleeve 60, and then pass outwardly opening 61,62, for rushing of the blade outer wall on the opposite side in chamber 19
Hit cooling.
Insertion sleeve 60 have impact opening 61,62 scope depend on insertion sleeve side wall to be airfoil pressure
Side or suction side.Although if being considered as to realize needed for the purpose of the present invention or if expectation, impact opening is on every side
Scope can change, but the scope in the hole on the pressure side gone up of insertion sleeve is typically larger than suction side.
In the diagram, impinging cooling zigzag opening is similarly positioned in the upstream sleeve part 65,66 of insertion sleeve, and inserts
The other downstream sleeve part 67,68 entering sleeve does not have hole.On the contrary, downstream part makes the coolant flow area in chamber 19
Reduce.For the insertion sleeve in leading edge chamber and return cavity, using the upstream portion of new impact opening, mixing is improved,
Impinging air jet velocity improves and the heat transfer of those parts of blade is finally more preferable.
Although already in connection be presently considered to be can put into practice and preferred embodiment invention has been described,
It is it should be appreciated that the present invention is not limited to the disclosed embodiments, but include in claims Instead, it is desirable to cover
Various remodeling in spirit and scope and equivalent arrangements.
Claims (19)
1. a kind of nozzle vane for gas-turbine unit, described nozzle vane includes：
Blade wall, described blade wall has inner wall surface and outer wall surface, and described inner wall surface and outer wall surface are equably each other
It is spaced apart, to limit the fluid passage for cooling medium；
Multiple chambeies, the plurality of chamber is arranged between described inner wall surface and described outer wall surface and is located at described fluid and leads to
In road；
Multiple impinging cooling sleeves, the plurality of impinging cooling sleeve be arranged in by described inner wall surface and described outer wall surface with
And in the discrete described chamber of inside wall member restriction；And
Multiple non-circular impact opening in each of the plurality of impinging cooling sleeve impinging cooling sleeve, described
The size and number of multiple non-circular impact openings be enough to accommodate the stream entering the described cooling medium within described fluid passage
Dynamic, wherein said non-circular impact opening has the protuberance of arranged radially, and described protuberance is from each non-circular impact opening
Center radiation setting to form corresponding axial fluid path；Wherein, the rib of described non-circular impact opening and described cooling
The respective surfaces that medium enters the described impinging cooling sleeve of described non-circular impact opening are located in identical plane.
2. nozzle vane according to claim 1 is it is characterised in that the plurality of not rounded in described impinging cooling sleeve
Shape impact opening includes largely star shaped fluid passage, and described largely star shaped fluid passage has from each non-circular impact opening
The radial arrays of the protuberance of the radial setting in center.
3. nozzle vane according to claim 1 is it is characterised in that the plurality of not rounded in described impinging cooling sleeve
Shape impacts the array of general uniform, the equidistant fluid passage from the radial setting in the center of each opening for the limited opening.
4. nozzle vane according to claim 1 is it is characterised in that the plurality of not rounded in described impinging cooling sleeve
Shape impacts opening substantially along the length arrangement of each insertion sleeve.
5. nozzle vane according to claim 1 is it is characterised in that the plurality of not rounded in described impinging cooling sleeve
Shape impact opening includes chevron structure.
6. nozzle vane according to claim 1 is it is characterised in that spray described in the described impact flow of described cooling medium
The described inner wall surface of mouth blade.
7. nozzle vane according to claim 1 is it is characterised in that the plurality of not rounded in described impinging cooling sleeve
Shape impact opening is formed in first wall and the second wall of each of multiple impinging cooling sleeves impinging cooling sleeve.
8. nozzle vane according to claim 1 it is characterised in that by described impinging cooling sleeve formed described discrete
Chamber have and described internal wall member identical total structure.
9. nozzle vane according to claim 1 is it is characterised in that described discrete chamber and described impinging cooling sleeve limit
Determine the specific fluid flow gap for described cooling medium.
10. nozzle vane according to claim 1 is it is characterised in that described discrete chamber longitudinal extension, and described
Parallel to each other in nozzle vane.
11. nozzle vanes according to claim 1 are it is characterised in that described cooling medium includes compressed air.
12. nozzle vanes according to claim 1 are it is characterised in that described cooling medium includes steam.
13. nozzle vanes according to claim 1 are it is characterised in that described cooling medium is steam and compressed air
Mixture.
14. nozzle vanes according to claim 1 are it is characterised in that the shape of the forward edge of described impinging cooling sleeve
Shape is bending, and the side wall of described impinging cooling sleeve is corresponding generally to the shape of the side wall in described chamber.
15. nozzle vanes according to claim 1 are it is characterised in that the plurality of non-in described impinging cooling sleeve
Circular impact opening makes the air mixing along described nozzle vane improve.
16. nozzle vanes according to claim 1 are it is characterised in that the plurality of non-in described impinging cooling sleeve
Circular impact opening makes the impinging air jet velocity along described nozzle vane improve.
17. nozzle vanes according to claim 1 are it is characterised in that the plurality of non-circular impact opening makes along institute
The crossing current heat transfer degradation stating nozzle vane reduces.
18. nozzle vanes according to claim 1 are it is characterised in that the plurality of non-circular impact opening is provided only on
In the upstream portion of described impinging cooling sleeve.
19. nozzle vanes according to claim 1 are it is characterised in that the plurality of non-circular impact opening has sawtooth
Shape is arranged.
Applications Claiming Priority (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US13/326,372 | 2011-12-15 | ||
| US13/326372 | 2011-12-15 | ||
| US13/326,372 US9151173B2 (en) | 2011-12-15 | 2011-12-15 | Use of multi-faceted impingement openings for increasing heat transfer characteristics on gas turbine components |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN103161513A CN103161513A (en) | 2013-06-19 |
| CN103161513B true CN103161513B (en) | 2017-03-01 |
Family
ID=47297013
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201210545238.3A Expired - Fee Related CN103161513B (en) | 2011-12-15 | 2012-12-14 | The improved nozzle vane for gas-turbine unit |
Country Status (5)
| Country | Link |
|---|---|
| US (1) | US9151173B2 (en) |
| EP (1) | EP2604800B1 (en) |
| JP (1) | JP2013124663A (en) |
| CN (1) | CN103161513B (en) |
| RU (1) | RU2012153930A (en) |
Families Citing this family (26)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US10100666B2 (en) * | 2013-03-29 | 2018-10-16 | General Electric Company | Hot gas path component for turbine system |
| US9849510B2 (en) | 2015-04-16 | 2017-12-26 | General Electric Company | Article and method of forming an article |
| US9976441B2 (en) | 2015-05-29 | 2018-05-22 | General Electric Company | Article, component, and method of forming an article |
| US9995151B2 (en) | 2015-08-17 | 2018-06-12 | General Electric Company | Article and manifold for thermal adjustment of a turbine component |
| US10087776B2 (en) | 2015-09-08 | 2018-10-02 | General Electric Company | Article and method of forming an article |
| US10253986B2 (en) * | 2015-09-08 | 2019-04-09 | General Electric Company | Article and method of forming an article |
| US10739087B2 (en) | 2015-09-08 | 2020-08-11 | General Electric Company | Article, component, and method of forming an article |
| US10815789B2 (en) | 2016-02-13 | 2020-10-27 | General Electric Company | Impingement holes for a turbine engine component |
| EP3236010A1 (en) * | 2016-04-21 | 2017-10-25 | Siemens Aktiengesellschaft | Stator vane having a junction tubing |
| US10309228B2 (en) * | 2016-06-09 | 2019-06-04 | General Electric Company | Impingement insert for a gas turbine engine |
| US10450868B2 (en) * | 2016-07-22 | 2019-10-22 | General Electric Company | Turbine rotor blade with coupon having corrugated surface(s) |
| US10465520B2 (en) | 2016-07-22 | 2019-11-05 | General Electric Company | Blade with corrugated outer surface(s) |
| US10465525B2 (en) | 2016-07-22 | 2019-11-05 | General Electric Company | Blade with internal rib having corrugated surface(s) |
| US10436037B2 (en) | 2016-07-22 | 2019-10-08 | General Electric Company | Blade with parallel corrugated surfaces on inner and outer surfaces |
| US10443399B2 (en) * | 2016-07-22 | 2019-10-15 | General Electric Company | Turbine vane with coupon having corrugated surface(s) |
| US10655477B2 (en) | 2016-07-26 | 2020-05-19 | General Electric Company | Turbine components and method for forming turbine components |
| US10577942B2 (en) * | 2016-11-17 | 2020-03-03 | General Electric Company | Double impingement slot cap assembly |
| EP3351341A1 (en) | 2017-01-23 | 2018-07-25 | Siemens Aktiengesellschaft | Method for producing a cavity in a blade platform |
| US20190277501A1 (en) * | 2018-03-07 | 2019-09-12 | United Technologies Corporation | Slot arrangements for an impingement floatwall film cooling of a turbine engine |
| US11192207B2 (en) | 2018-10-26 | 2021-12-07 | General Electric Company | Additive manufactured object with passage having varying cross-sectional shape |
| US10711620B1 (en) * | 2019-01-14 | 2020-07-14 | General Electric Company | Insert system for an airfoil and method of installing same |
| US11358221B2 (en) * | 2019-09-30 | 2022-06-14 | The Boeing Company | Build part and method of additively manufacturing the build part |
| US11629642B2 (en) | 2019-12-20 | 2023-04-18 | General Electric Company | System and methods for igniting and operating a gas turbine engine with alternative fuels |
| US11959641B2 (en) * | 2020-01-31 | 2024-04-16 | Rtx Corporation | Combustor shell with shaped impingement holes |
| CN112268012B (en) * | 2020-10-10 | 2022-02-11 | 浙江理工大学 | Spiralless centrifugal fan impeller with tail jet device and its working method |
| CN113944516B (en) * | 2021-09-28 | 2024-04-02 | 中国科学院工程热物理研究所 | Composite cooling structure for tip of gas turbine |
Family Cites Families (39)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US1545560A (en) | 1924-08-06 | 1925-07-14 | Heath Spencer | Airplane propeller |
| US1939357A (en) | 1929-02-13 | 1933-12-12 | Bendix Aviat Corp | Hollow blade for turbines |
| US3148954A (en) | 1960-06-13 | 1964-09-15 | Haas Irene | Turbine blade construction |
| GB1565361A (en) | 1976-01-29 | 1980-04-16 | Rolls Royce | Blade or vane for a gas turbine engien |
| GB2119028B (en) | 1982-04-27 | 1985-02-27 | Rolls Royce | Aerofoil for a gas turbine engine |
| US4645415A (en) | 1983-12-23 | 1987-02-24 | United Technologies Corporation | Air cooler for providing buffer air to a bearing compartment |
| FR2581708B1 (en) | 1985-05-09 | 1989-04-28 | Snecma | COVER FOR TURBOREACTOR BLADE BLADE ATTACK EDGE |
| US4669957A (en) | 1985-12-23 | 1987-06-02 | United Technologies Corporation | Film coolant passage with swirl diffuser |
| FR2599425B1 (en) | 1986-05-28 | 1988-08-05 | Alsthom | PROTECTIVE PLATE FOR TITANIUM BLADE AND METHOD OF BRAZING SUCH A PLATE. |
| US5667359A (en) | 1988-08-24 | 1997-09-16 | United Technologies Corp. | Clearance control for the turbine of a gas turbine engine |
| US5210946A (en) | 1992-06-26 | 1993-05-18 | Hudson Products Corporation | Leading edge protection for fan blade |
| JP3651490B2 (en) | 1993-12-28 | 2005-05-25 | 株式会社東芝 | Turbine cooling blade |
| US5634766A (en) * | 1994-08-23 | 1997-06-03 | General Electric Co. | Turbine stator vane segments having combined air and steam cooling circuits |
| GB2293631B (en) | 1994-09-30 | 1998-09-09 | Gen Electric | Composite fan blade trailing edge reinforcement |
| US5685693A (en) | 1995-03-31 | 1997-11-11 | General Electric Co. | Removable inner turbine shell with bucket tip clearance control |
| US5779437A (en) | 1996-10-31 | 1998-07-14 | Pratt & Whitney Canada Inc. | Cooling passages for airfoil leading edge |
| JPH1144498A (en) * | 1997-05-30 | 1999-02-16 | Showa Alum Corp | Flat perforated tube for heat exchanger and heat exchanger using the tube |
| US6238182B1 (en) | 1999-02-19 | 2001-05-29 | Meyer Tool, Inc. | Joint for a turbine component |
| US6183192B1 (en) | 1999-03-22 | 2001-02-06 | General Electric Company | Durable turbine nozzle |
| JP3794868B2 (en) | 1999-06-15 | 2006-07-12 | 三菱重工業株式会社 | Gas turbine stationary blade |
| JP3322857B2 (en) | 1999-11-30 | 2002-09-09 | モルデック株式会社 | Insert molding method for receptacle connector |
| US6468031B1 (en) * | 2000-05-16 | 2002-10-22 | General Electric Company | Nozzle cavity impingement/area reduction insert |
| US6428273B1 (en) * | 2001-01-05 | 2002-08-06 | General Electric Company | Truncated rib turbine nozzle |
| JP4508432B2 (en) | 2001-01-09 | 2010-07-21 | 三菱重工業株式会社 | Gas turbine cooling structure |
| DE50208671D1 (en) * | 2001-07-13 | 2006-12-21 | Alstom Technology Ltd | Gas turbine section with cooling air hole |
| US6742991B2 (en) | 2002-07-11 | 2004-06-01 | Mitsubishi Heavy Industries, Ltd. | Turbine blade and gas turbine |
| JP4191578B2 (en) | 2003-11-21 | 2008-12-03 | 三菱重工業株式会社 | Turbine cooling blade of gas turbine engine |
| US7008178B2 (en) * | 2003-12-17 | 2006-03-07 | General Electric Company | Inboard cooled nozzle doublet |
| US20050265840A1 (en) | 2004-05-27 | 2005-12-01 | Levine Jeffrey R | Cooled rotor blade with leading edge impingement cooling |
| US7374401B2 (en) * | 2005-03-01 | 2008-05-20 | General Electric Company | Bell-shaped fan cooling holes for turbine airfoil |
| US20100143154A1 (en) * | 2007-01-04 | 2010-06-10 | Ansaldo Energia S.P.A. | Spacer for gas turbine blade insert |
| DE102007018061A1 (en) * | 2007-04-17 | 2008-10-23 | Rolls-Royce Deutschland Ltd & Co Kg | Gas turbine combustion chamber wall |
| JP2009162119A (en) * | 2008-01-08 | 2009-07-23 | Ihi Corp | Turbine blade cooling structure |
| US20100037620A1 (en) | 2008-08-15 | 2010-02-18 | General Electric Company, Schenectady | Impingement and effusion cooled combustor component |
| US8142137B2 (en) * | 2008-11-26 | 2012-03-27 | Alstom Technology Ltd | Cooled gas turbine vane assembly |
| CH701031A1 (en) * | 2009-05-15 | 2010-11-15 | Alstom Technology Ltd | The method for refurbishing a turbine blade. |
| US20110097191A1 (en) * | 2009-10-28 | 2011-04-28 | General Electric Company | Method and structure for cooling airfoil surfaces using asymmetric chevron film holes |
| US8529193B2 (en) * | 2009-11-25 | 2013-09-10 | Honeywell International Inc. | Gas turbine engine components with improved film cooling |
| FR2955145B1 (en) * | 2010-01-14 | 2012-02-03 | Snecma | HIGH PRESSURE TURBINE DISPENSER OF A TURBOREACTOR |
-
2011
- 2011-12-15 US US13/326,372 patent/US9151173B2/en not_active Expired - Fee Related
-
2012
- 2012-12-07 EP EP12196112.2A patent/EP2604800B1/en not_active Not-in-force
- 2012-12-10 JP JP2012268914A patent/JP2013124663A/en active Pending
- 2012-12-14 CN CN201210545238.3A patent/CN103161513B/en not_active Expired - Fee Related
- 2012-12-14 RU RU2012153930/06A patent/RU2012153930A/en not_active Application Discontinuation
Also Published As
| Publication number | Publication date |
|---|---|
| EP2604800B1 (en) | 2018-02-21 |
| EP2604800A2 (en) | 2013-06-19 |
| RU2012153930A (en) | 2014-06-20 |
| US9151173B2 (en) | 2015-10-06 |
| EP2604800A3 (en) | 2015-07-22 |
| JP2013124663A (en) | 2013-06-24 |
| US20130156549A1 (en) | 2013-06-20 |
| CN103161513A (en) | 2013-06-19 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN103161513B (en) | The improved nozzle vane for gas-turbine unit | |
| US7255534B2 (en) | Gas turbine vane with integral cooling system | |
| CN101852099B (en) | Film-cooling augmentation device and turbine airfoil incorporating the same | |
| US7390168B2 (en) | Vortex cooling for turbine blades | |
| CN100385091C (en) | High temperature gas flow passage component with mesh type and impingement cooling | |
| US8721281B2 (en) | Cooled blade for a gas turbine | |
| CA2642505C (en) | Cooling structure | |
| JP6931992B2 (en) | Gradual fuel and air injection in the combustion system of a gas turbine | |
| US7513739B2 (en) | Cooling circuits for a turbomachine moving blade | |
| US9347324B2 (en) | Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles | |
| US20100221121A1 (en) | Turbine airfoil cooling system with near wall pin fin cooling chambers | |
| US8801366B2 (en) | Stator blade for a gas turbine and gas turbine having same | |
| US7597540B1 (en) | Turbine blade with showerhead film cooling holes | |
| US8511968B2 (en) | Turbine vane for a gas turbine engine having serpentine cooling channels with internal flow blockers | |
| CA2528098C (en) | Internally cooled gas turbine airfoil and method | |
| CN106133276B (en) | Turbine airfoil | |
| CN106795771A (en) | Inner cooling system with the insert that nearly wall cooling duct is formed in cooling chamber in the middle part of the wing chord of gas turbine aerofoil profile | |
| US7287955B2 (en) | Gas turbine clearance control devices | |
| US20120070302A1 (en) | Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles | |
| CN101886553A (en) | Turbine nozzle with sidewall cooling plenum | |
| CN106661945A (en) | Internal Cooling System With Insert Forming Nearwall Cooling Channels In An Aft Cooling Cavity Of A Gas Turbine Airfoil | |
| CN106471212A (en) | There is leading edge impinging cooling system and the turbine airfoil cooling system of nearly wall impact system | |
| JP6996947B2 (en) | Turbine blades and gas turbines | |
| US11346248B2 (en) | Turbine nozzle segment and a turbine nozzle comprising such a turbine nozzle segment | |
| KR20250111875A (en) | Turbine blade and gas turbine comprising the same |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C06 | Publication | ||
| PB01 | Publication | ||
| C10 | Entry into substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant | ||
| CF01 | Termination of patent right due to non-payment of annual fee | ||
| CF01 | Termination of patent right due to non-payment of annual fee |
Granted publication date: 20170301 Termination date: 20181214 |